CN116357404A - Braking structure for preventing turbine from flying after broken shaft of aero-engine - Google Patents
Braking structure for preventing turbine from flying after broken shaft of aero-engine Download PDFInfo
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- CN116357404A CN116357404A CN202310560800.8A CN202310560800A CN116357404A CN 116357404 A CN116357404 A CN 116357404A CN 202310560800 A CN202310560800 A CN 202310560800A CN 116357404 A CN116357404 A CN 116357404A
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- pressure turbine
- conical surface
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/16—Arrangement of bearings; Supporting or mounting bearings in casings
- F01D25/162—Bearing supports
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention belongs to the field of safety design of aeroengines, and discloses a brake structure for preventing a turbine from flying after an aeroengine breaks a shaft, wherein the brake structure is positioned between a low-pressure turbine and a turbine rear bearing frame, the rear side of a turbine rear journal is fixedly connected with a short drum, the short drum is annular, and the radial inner side of the short drum is distributed with the low-pressure turbine shaft at intervals and is coaxially arranged with the low-pressure turbine shaft; in the axial direction, the gap between the rear end of the short drum and the rigid bearing seat is distributed; when the low-pressure turbine shaft is broken, the low-pressure turbine moves backwards, friction occurs between the rear end of the rear shaft neck of the turbine and the rigid bearing seat, the rotation speed of the low-pressure turbine is limited, and the low-pressure turbine is braked.
Description
Technical Field
The invention belongs to the field of safety design of aeroengines, and particularly relates to a brake structure for preventing a turbine from flying after an aeroengine breaks a shaft.
Background
Turbine turning is a severe and complex accident for aircraft engines. During normal operation of an aeroengine, because a low-pressure turbine shaft breaks due to extreme load, machining and assembling, fatigue, clamping stagnation or other factors, after the low-pressure turbine shaft breaks, a low-pressure rotor loses load, the low-pressure turbine speed is increased under the driving of fuel gas, when the turbine speed is increased to a certain level, turbine blades break, an impact casing is impacted, a turbine disc breaks due to overlarge centrifugal stress, and generated high-energy fragments can break through the casing to fly outwards, possibly hit an airplane, and the safety of the airplane is seriously affected. Therefore, designing a brake structure for preventing the turbine from flying after the aero-engine is an indispensable content in the safety design.
In order to reduce the damage caused by turbine turning, the traditional method is to actively control by a method of turning fuel cut off and turning down through rotor rotation speed monitoring, but the method has the defects that the monitoring equipment stops fuel after finding that an engine rotor turns, fuel is stopped, fuel is still supplied in an oil pipe at the moment of stopping fuel, and the engine rotor still continues to rotate, so that the rotation speed of an engine rotor system is faster and faster, and an irrecoverable secondary accident occurs. Stopping the fuel does not immediately cause the rotor system to lose drive, and this process is time-delayed. With the development of technology, the active control braking technology is no longer applicable to modern aeroengines, so that a braking structure needs to be designed to prevent the turbine from flying by a passive control method.
After the low-pressure rotor of the aeroengine breaks the shaft, the low-pressure turbine can impact the bearing frame backwards under the driving of fuel gas, but the bearing frame is elastic, certain rebound can occur after the turbine rotor impacts the bearing frame, the rebound can cause the collision friction between the low-pressure turbine and other stator structures after the turbine rotor impacts the bearing frame, and under the condition of high rotating speed, secondary accidents can be caused, so that the structural integrity of the aeroengine is damaged, and therefore, the design of a rebound limiting structure after the shaft breaking is necessary for the aeroengine so as to improve the safety of the aeroengine.
Disclosure of Invention
In order to solve the technical problems, the invention provides a brake structure for preventing a turbine from flying after an aeroengine breaks a shaft, so as to solve the problems in the prior art, and in order to achieve the purposes of the invention, the invention adopts the following technical scheme:
the brake structure is used for preventing a turbine from flying after an aeroengine breaks a shaft and is applied to the aeroengine, the brake structure is positioned between a low-pressure turbine and a turbine rear bearing frame, a turbine rear journal is fixedly arranged at the rear end of the low-pressure turbine, the turbine rear journal is connected with a low-pressure turbine shaft, a fulcrum bearing is arranged on the low-pressure turbine shaft, and the fulcrum bearing is connected with the turbine rear bearing frame through a rigid bearing seat;
the rear side of the rear shaft neck of the turbine is fixedly connected with a short drum barrel, and the short drum barrel is annular, is distributed between the radial inner side of the short drum barrel and the low-pressure turbine shaft at intervals and is coaxially arranged with the low-pressure turbine shaft; in the axial direction, the gap between the rear end of the short drum and the rigid bearing seat is distributed;
when the low-pressure turbine shaft breaks, the low-pressure turbine moves backwards, friction occurs between the rear end of the rear shaft neck of the turbine and the rigid bearing seat, and the rotation speed of the low-pressure turbine is limited and the low-pressure turbine is braked.
Further, the rear end of the short drum is provided with a wedge-shaped conical surface, the rigid bearing seat comprises a front edge conical surface which is matched with the wedge-shaped conical surface and corresponds to the wedge-shaped conical surface in position, and the wedge-shaped conical surface and the front edge conical surface are of conical surface structures and are in gap distribution.
Further, a first friction part is arranged in the circumferential direction of the wedge-shaped conical surface, a second friction part is arranged in the circumferential direction of the front edge conical surface correspondingly, and the first friction part and the second friction part are used for improving friction force during contact.
Further, the first friction part is a first convex tooth fixedly arranged on the wedge-shaped conical surface, and the second friction part is a second convex tooth fixedly arranged on the front edge conical surface.
Further, the rigid bearing seat further comprises an axial flange arranged at the front end of the front edge cone, a ring groove matched with the axial flange is arranged at the radial outer side of the short drum, and a lug is fixedly arranged at the radial inner side of the rear end of the short drum;
when the low-pressure turbine shaft breaks, the rotation speed of the low-pressure turbine is increased, the short drum barrel is radially deformed, the radial interval between the short drum barrel and the axial flange at the front end of the rigid bearing seat is reduced, the annular grooves are nested on the axial flange and are mutually clamped, so that the low-pressure turbine is axially limited, the rebound limiting function of the low-pressure turbine is realized, and the protruding blocks are used for increasing the radial deformation of the short drum barrel.
Further, the rigid bearing seat further comprises a bearing installation part, a force transmission conical shell and a fixed installation edge which are sequentially arranged, the front edge conical surface is positioned at the front end of the bearing installation part, the bearing outer ring of the fulcrum bearing is installed on the radial inner side of the bearing installation part, the force transmission conical shell is obliquely arranged, and the rear end of the fixed installation edge is connected with the turbine rear bearing frame.
Further, the turbine rear journal comprises a spigot mounting edge and a journal, the rear side of the journal is fixedly connected with the front end of the short drum, the radial outer side of the journal is fixedly connected with the spigot mounting edge, and the spigot mounting edge is detachably connected with the low-pressure turbine.
Further, the radial inner side of the journal is connected with the low-pressure turbine shaft through a sleeve gear coupling.
Further, a groove body is formed between the radial inner side of the shaft neck and the radial outer side of the low-pressure turbine shaft, the tooth-sleeved coupler is positioned in the groove body, the rear end of the shaft neck is abutted against the front end face of the fulcrum bearing, the rear end face of the fulcrum bearing is abutted against a nut, and the nut is sleeved on the low-pressure turbine shaft.
The utility model provides an aeroengine, this aeroengine includes three-stage fan, intermediary load frame, five high-pressure compressor, annular combustion chamber, one-level high-pressure turbine, turbine interstage load frame, low-pressure turbine and turbine back load frame that from front to back set gradually, and this brake structure is located between low-pressure turbine and the turbine back load frame, the low-pressure turbine is two-stage low-pressure turbine.
The invention has the following beneficial effects: according to the invention, the rapid braking and rebound limiting of the turbine flywheel of the aeroengine are realized by adopting the wedge-shaped conical surface of the rear shaft neck of the turbine and the conical surface of the front edge of the bearing seat. When the aeroengine normally operates, a gap exists between the wedge-shaped conical surface of the rear shaft diameter of the turbine and the front edge of the bearing seat, and contact friction cannot occur. When the aeroengine breaks the shaft, the turbine rotor moves backwards, the wedge-shaped conical surface of the rear shaft neck of the turbine rubs with the conical surface of the front edge of the bearing seat, the rotating speed of the turbine rotor is limited through friction, and the turbine rotor is rapidly braked. The turbine can rebound after impacting the rigid bearing seat backwards, the rebound limit of the turbine is realized through the axial flange and the annular groove, the structural integrity of the turbine component is ensured, the occurrence of secondary accidents is prevented, and the safety is higher.
Drawings
FIG. 1 is a schematic diagram of a typical dual rotor high thrust-to-weight ratio turbofan aircraft engine of the present invention.
FIG. 2 is a partial block diagram of the low pressure turbine and turbine aft thrust frame of FIG. 1;
FIG. 3 is an enlarged view of a portion of FIG. 2 at A;
FIG. 4 is an enlarged view of the brake structure at B in FIG. 3;
FIG. 5 is an annular cutaway view at C in FIG. 4;
FIG. 6 is an enlarged view of the braking structure after turbine turning in the present invention.
Fig. 7 is an annular cutaway view at D in fig. 6.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to fig. 1 to 7 in the embodiments of the present invention, and it is obvious that the described embodiments are only some embodiments of the present invention, but not all embodiments, and technical means used in the embodiments are conventional means known to those skilled in the art unless specifically indicated.
1-7 are schematic cross-sectional views, in which the low pressure turbine shaft 90, the set tooth coupling 2, the turbine rear journal 1, the low pressure turbine 97 and the fulcrum bearing 4 are all coaxial rotary structures, which are referred to herein simply as a low pressure turbine rotor system; the rigid bearing seat 3 and the turbine rear bearing frame 98 are of a coaxial non-rotating structure, and are called a stator system; arrows in fig. 1 and 6 are the moving direction of the low-pressure turbine 97 after the low-pressure turbine shaft 90 breaks.
In addition, the "radial direction" referred to in the present invention is the up-down direction in the drawing, and the "radial outside" means the position where the member is located above, that is, the member is located on the side away from the low-pressure turbine shaft 90; "radially inward" means that the component is located in a lower position, i.e., the component is located on a side adjacent to the low pressure turbine shaft 90; the term "front end" refers to the axial left side in the drawing, the term "rear end" refers to the axial right side in the drawing, and the terms "front end" and "rear end" are defined by the intake and exhaust directions of the engine, and the front end of the engine is intake and the rear end of the engine is exhaust according to industry conventions.
Referring to fig. 1, a typical dual-rotor high thrust-weight ratio turbofan aeroengine 9 is specifically configured, the aeroengine 9 includes three stages of fans 91, an intermediate bearing frame 92, five stages of high-pressure compressors 93, an annular combustion chamber 94, a first stage high-pressure turbine 95, a turbine inter-stage bearing frame 96, a low-pressure turbine 97 and a turbine rear bearing frame 98, which are sequentially arranged from front to back, and the brake structure is located between the low-pressure turbine 97 and the turbine rear bearing frame 98, and the low-pressure turbine 97 is a two-stage low-pressure turbine.
The present invention relates to a brake structure for preventing turbine from flying, which is designed between a low-pressure turbine 97 and a turbine rear bearing frame 98, that is, the present invention mainly aims at the brake structure for preventing turbine from flying for two-stage low-pressure turbine.
As shown in fig. 2, a brake structure for preventing a turbine from flying after an aeroengine breaks a shaft is applied to an aeroengine 9, the brake structure is positioned between a low-pressure turbine 97 and a turbine rear bearing frame 98, a turbine rear journal 1 is fixedly arranged at the rear end of the low-pressure turbine 97, the turbine rear journal 1 is connected with a low-pressure turbine shaft 90, a fulcrum bearing 4 is arranged on the low-pressure turbine shaft 90, and the fulcrum bearing 4 is connected with the turbine rear bearing frame 98 through a rigid bearing seat 3;
the rear side of the turbine rear journal 1 is fixedly connected with a short drum 13, the short drum 13 is annular, and the radial inner side of the short drum 13 is distributed with the low-pressure turbine shaft 90 at intervals and is coaxially arranged with the low-pressure turbine shaft 90; in the axial direction, the gap between the rear end of the short drum 13 and the rigid bearing seat 3 is distributed;
when the low pressure turbine shaft 90 breaks, the low pressure turbine 97 moves backward, friction occurs between the rear end of the turbine rear journal 1 and the rigid bearing housing 3, and the rotation speed of the low pressure turbine 97 is limited and braked.
The low pressure turbine 97 and the turbine aft bearing frame 98 are both prior art structures of the aircraft engine 9, the rotating component turbine aft journal 1 and the non-rotating component rigid bearing mount 3 are located between the low pressure turbine 97 and the turbine aft bearing frame 98, and the turbine aft journal 1 is located at the aft end of the low pressure turbine 97, with the rigid bearing mount 3 being located at a position between the radially inner side of the turbine aft bearing frame 98 and the low pressure turbine shaft 90.
As shown in fig. 6 and 7, a reserved gap is designed between the rear end of the short drum 13 and the rigid bearing seat 3, when the low-pressure turbine shaft 90 breaks, the rotating part low-pressure turbine 97, the turbine rear journal 1 and the low-pressure turbine shaft 90 can move backwards under the action of pneumatic force, and because of losing load, the rotating speed can be faster and faster under the driving of fuel gas, and at the moment, the short drum 13 is contacted with the rigid bearing seat 3 to generate sliding friction, so that the purpose of decelerating and realizing quick braking is achieved.
Referring to fig. 3 and 4, the turbine rear journal 1 integrally comprises an integrally formed spigot mounting edge 11, a journal 12, a short drum 13 and a wedge-shaped conical surface 14, the wedge-shaped conical surface 14 is arranged at the rear end of the short drum 13, the rigid bearing seat 3 comprises a front edge conical surface 32 which is matched with the wedge-shaped conical surface 14 and corresponds to the wedge-shaped conical surface in position, the wedge-shaped conical surface 14 and the front edge conical surface 32 are of conical surface structures, and gaps between the wedge-shaped conical surface and the front edge conical surface 32 are distributed.
Specifically, the wedge-shaped conical surface 14 is an annular conical surface, the outer diameter of the wedge-shaped conical surface is gradually shortened from front to back, the front edge conical surface 32 is also an annular structure, the radial interval between the wedge-shaped conical surface 14 and the low-pressure turbine shaft 90 is gradually shortened from front to back, and after the low-pressure turbine shaft 90 is broken, the wedge-shaped conical surface 14 and the front edge conical surface 32 can generate contact friction to realize rapid braking.
Further, a first friction portion is disposed in the circumferential direction of the wedge-shaped tapered surface 14, and a second friction portion is disposed in the circumferential direction of the leading edge tapered surface 32, and the first friction portion and the second friction portion are used for improving friction force during contact.
Further, the first friction portion is a first tooth 141 fixedly disposed on the wedge-shaped conical surface 14, and the second friction portion is a second tooth 321 fixedly disposed on the leading edge conical surface 32.
The first convex teeth and the second convex teeth are distributed annularly and uniformly, the first convex teeth and the second convex teeth are used for improving sliding friction force and are meshed to form a certain rotation limit, and the rotor system can be braked rapidly.
Further, the rigid bearing seat 3 further includes an axial flange 31 disposed at the front end of the front edge conical surface 32, a ring groove 131 adapted to the axial flange 31 is disposed at the radial outer side of the short drum 13, and a protruding block 142 is fixedly disposed at the radial inner side of the rear end of the short drum 13;
when the low-pressure turbine shaft 90 breaks, the rotation speed of the low-pressure turbine 97 increases, the short drum 13 deforms radially, the radial interval between the short drum and the axial flange 31 at the front end of the rigid bearing seat 3 decreases, the annular grooves 131 are nested on the axial flange 31 and are clamped mutually, so that the low-pressure turbine 97 is axially limited, the rebound limiting function of the low-pressure turbine 97 is realized, and the protruding blocks 142 are used for increasing the radial deformation of the short drum 13.
The axial rib 31 forms a convex structure at the front end of the front edge conical surface 32, the annular groove 131 is radially opposite to the axial rib 31, and a certain clearance exists, and the radial deformation of the short drum 13 is smaller than the clearance during normal operation of the engine; after the low-pressure turbine shaft 90 is broken, the turbine rotor moves backwards under the action of pneumatic force and flies, the turbine rotor rebounds after impacting the rigid bearing seat 3, the short drum 13 deforms and increases in radial direction due to the flying, the annular groove 131 and the axial flange 31 are clamped mutually, and the rebound limit of the low-pressure turbine is realized.
Referring to fig. 3 and 4, the rigid bearing seat 3 integrally includes an integrally formed axial flange 31, a front edge conical surface 32, a bearing mounting portion 33, a force transmission conical shell 34 and a fixed mounting edge 35, the front edge conical surface 32 is located at the front end of the bearing mounting portion 33, the bearing outer ring 41 of the fulcrum bearing 4 is mounted on the radial inner side of the bearing mounting portion 33, the force transmission conical shell 34 is obliquely arranged, and the rear end of the fixed mounting edge 35 is connected with the turbine rear bearing frame 98.
Specifically, the bearing mounting portion 33 is parallel to the low pressure turbine shaft 90, the force transmitting cone shell 34 is disposed obliquely, and the radially outer side of the fixed mounting edge 35 is bolted to the mounting edge 981 of the turbine aft load frame 98.
Further, the rear side of the journal 12 is fixedly connected to the front end of the short drum 13, the radially outer side of the journal 12 is fixedly connected to the spigot mounting edge 11, and the spigot mounting edge 11 is detachably connected to the low pressure turbine 97.
Specifically, the spigot mounting edge 11 radially outside the turbine rear journal 1 is connected to the flange edge 971 of the low pressure turbine 97 and the mounting edge 51 of the seal structure 5 by bolts 6.
Further, the radially inner side of the journal 12 is connected to the low pressure turbine shaft 90 through a set-tooth coupling 2.
Further, a groove body is formed between the radial inner side of the journal 12 and the radial outer side of the low-pressure turbine shaft 90, the set gear coupling 2 is located in the groove body, the rear end of the journal 12 abuts against the front end face of the fulcrum bearing 4, the rear end face of the fulcrum bearing 4 abuts against the nut 8, and the nut 8 is sleeved on the low-pressure turbine shaft 90.
Specifically, the bearing 4 comprises a bearing outer ring 41, rollers 42, a retainer 43 and a bearing inner ring 44, the bearing outer ring 41 abuts against the inner side of the bearing mounting part 33, and positioning mounting of the bearing is realized through the bearing flange 7; the rear end of the bearing inner ring 44 is axially positioned by the nut 8, the front end of the bearing inner ring 44 abuts the rear end of the journal 12, and rotates with the low pressure turbine shaft 90,
the journal 12 is in an L-shaped structure, the inner side edge of the journal is long, the journal is in a stepped structure, the radial outer side of the low-pressure turbine shaft 90 is provided with a corresponding stepped part, the tooth sleeving coupler 2 comprises a first radial positioning surface 21, a second radial positioning surface 24, an axial positioning surface 23 and a tooth sleeving 22, the first radial positioning surface 21 and the second radial positioning surface 24 are used for guaranteeing coaxiality of the turbine rear journal 1 and the low-pressure turbine shaft 90, the axial positioning surface 23 is used for axially positioning and mounting the turbine rear journal 1, and the tooth sleeving 22 transmits torque through mutual engagement. The sleeve teeth 22 transmit the torque of the turbine to the compressor, the compressor consumes energy, after the shaft is broken, the energy consumption component is lost, and the rotation speed of the turbine continuously rises under the blowing of the fuel gas, namely under the action of pneumatic force, so that the turbine flies.
The working principle of the braking structure for preventing the turbine from flying in the invention is described as follows:
the wedge-shaped conical surface 14 and the front edge conical surface 32 are opposite to each other in front-back direction and have a certain clearance, contact cannot occur during normal operation, the axial clearance is smaller than the axial clearance between the turbine rotor and the turbine stator, and when the low-pressure shaft 90 of the engine 9 breaks, the low-pressure turbine rotor moves backwards, and the wedge-shaped conical surface 14 is in initial contact with the front conical surface 32; the wedge-shaped conical surface 14 is provided with a first convex tooth 141, the front edge conical surface 32 is provided with a second convex tooth 321, the thicknesses of the first convex tooth 141 and the second convex tooth 321 are only 0.2-0.3mm, and the wedge-shaped conical surface is mainly used for increasing the friction force when the first convex tooth 141 and the second convex tooth 321 are contacted with each other so as to realize the rapid braking of the low-pressure turbine rotor.
The short drum 13 comprises a through hole 132 and a ring groove 131 which are uniformly distributed in the circumferential direction; the purpose of the through holes 132 is to prevent the oil from concentrating on the inner annular surface of the short drum 13, the annular groove 132 being radially opposite the axial flange 31 with a clearance, less than which the radial deformation of the short drum 13 occurs during normal engine operation; after the low-pressure turbine shaft 90 is broken, the turbine rotor moves backwards under the action of pneumatic force and flies, the turbine rotor rebounds after impacting the rigid bearing seat 3, the short drum 13 deforms radially and increases because of the rising of the rotating speed, and the annular groove 131 and the axial flange 31 are clamped mutually, so that the rebound limit of the low-pressure turbine is realized.
The wedge-shaped conical surface 14 is opposite to the front edge conical surface 32, when the engine works normally, contact does not occur, and after the low-pressure turbine breaks the shaft and rotates, the wedge-shaped conical surface 14 contacts with the front edge conical surface 32 to rub so as to realize the rapid braking of the low-pressure turbine rotor. Compared with other braking structures positioned at the blade tip and the disk edge of the turbine, the braking structure between the rear shaft journal 1 of the turbine and the rigid bearing 3 has smaller radius, higher coaxiality and higher stability, and can properly reduce the radial amplitude of the turbine flywheel during braking, thereby ensuring the stable realization of the rapid braking of the flywheel.
The above embodiments are only illustrative of the preferred embodiments of the present invention and are not intended to limit the scope of the present invention, and various modifications, variations, alterations, substitutions made by those skilled in the art to the technical solution of the present invention should fall within the protection scope defined by the claims of the present invention without departing from the spirit of the design of the present invention.
Claims (10)
1. A brake structure for preventing turbine from flying after aeroengine broken axle is applied on aeroengine (9), its characterized in that: the brake structure is positioned between a low-pressure turbine (97) and a turbine rear bearing frame (98), a turbine rear journal (1) is fixedly arranged at the rear end of the low-pressure turbine (97), the turbine rear journal (1) is connected with a low-pressure turbine shaft (90), a fulcrum bearing (4) is arranged on the low-pressure turbine shaft (90), and the fulcrum bearing (4) is connected with the turbine rear bearing frame (98) through a rigid bearing seat (3);
the rear side of the turbine rear journal (1) is fixedly connected with a short drum (13), the short drum (13) is annular, and the radial inner side of the short drum is distributed with the low-pressure turbine shaft (90) at intervals and is coaxially arranged with the low-pressure turbine shaft (90); in the axial direction, the gap between the rear end of the short drum (13) and the rigid bearing seat (3) is distributed;
when the low-pressure turbine shaft (90) breaks, the low-pressure turbine (97) moves backwards, friction occurs between the rear end of the turbine rear journal (1) and the rigid bearing seat (3), and the rotation speed of the low-pressure turbine (97) is limited and braked.
2. A brake structure for preventing turbine spin after an aero-engine shaft break according to claim 1, wherein: the rear end of the short drum (13) is provided with a wedge-shaped conical surface (14), the rigid bearing seat (3) comprises a front edge conical surface (32) which is matched with the wedge-shaped conical surface (14) and corresponds to the wedge-shaped conical surface in position, and the wedge-shaped conical surface (14) and the front edge conical surface (32) are of conical surface structures and are distributed in gaps.
3. A brake structure for preventing turbine spin after an aero-engine shaft break according to claim 2, wherein: a first friction part is arranged in the circumferential direction of the wedge-shaped conical surface (14), a second friction part is arranged in the circumferential direction of the front edge conical surface (32), and the first friction part and the second friction part are used for improving friction force during contact.
4. A brake structure for preventing turbine spin after an aero-engine shaft break according to claim 2, wherein: the first friction part is a first convex tooth (141) fixedly arranged on the wedge-shaped conical surface (14), and the second friction part is a second convex tooth (321) fixedly arranged on the front edge conical surface (32).
5. A brake structure for preventing turbine spin after an aero-engine shaft break according to claim 2, wherein: the rigid bearing seat (3) further comprises an axial flange (31) arranged at the front end of the front edge conical surface (32), a ring groove (131) matched with the axial flange (31) is arranged on the radial outer side of the short drum (13), and a lug (142) is fixedly arranged on the radial inner side of the rear end of the short drum (13);
when the low-pressure turbine shaft (90) breaks, the rotating speed of the low-pressure turbine (97) is increased, the short drum (13) is radially deformed, the radial interval between the short drum and the axial flange (31) at the front end of the rigid bearing seat (3) is reduced, the annular grooves (131) are nested on the axial flange (31) and are mutually clamped, so that the low-pressure turbine (97) is axially limited, the rebound limiting function of the low-pressure turbine (97) is realized, and the protruding blocks (142) are used for increasing the radial deformation of the short drum (13).
6. A brake structure for preventing turbine spin after an aero-engine shaft break according to claim 2, wherein: the rigid bearing seat (3) further comprises a bearing installation part (33), a force transmission conical shell (34) and a fixed installation edge (35) which are sequentially arranged, the front edge conical surface (32) is positioned at the front end of the bearing installation part (33), the radial inner side of the bearing installation part (33) is provided with a bearing outer ring (41) of the fulcrum bearing (4), the force transmission conical shell (34) is obliquely arranged, and the rear end of the fixed installation edge (35) is connected with a turbine rear bearing frame (98).
7. A brake structure for preventing turbine spin after an aero-engine shaft break according to claim 1, wherein: the turbine rear journal (1) comprises a spigot mounting edge (11) and a journal (12), the rear side of the journal (12) is fixedly connected with the front end of the short drum (13), the radial outer side of the journal (12) is fixedly connected with the spigot mounting edge (11), and the spigot mounting edge (11) is detachably connected with the low-pressure turbine (97).
8. The brake structure for preventing turbine spin after an aero-engine shaft is broken according to claim 7, wherein: the radial inner side of the journal (12) is connected with the low-pressure turbine shaft (90) through a sleeve gear coupling (2).
9. The brake structure for preventing turbine from flying after an aero-engine broken shaft according to claim 8, wherein: a groove body is formed between the radial inner side of the journal (12) and the radial outer side of the low-pressure turbine shaft (90), the tooth-sleeved coupler (2) is positioned in the groove body, the rear end of the journal (12) is abutted against the front end face of the fulcrum bearing (4), the rear end face of the fulcrum bearing (4) is abutted against the nut (8), and the nut (8) is sleeved on the low-pressure turbine shaft (90).
10. An aeroengine employing a braking structure according to any of claims 1-9, characterized in that: the aeroengine (9) comprises a three-stage fan (91), an intermediate bearing frame (92), a five-stage high-pressure compressor (93), an annular combustion chamber (94), a one-stage high-pressure turbine (95), a turbine inter-stage bearing frame (96), a low-pressure turbine (97) and a turbine rear bearing frame (98), wherein the three-stage fan (91), the intermediate bearing frame (92), the five-stage high-pressure compressor (93) and the annular combustion chamber are sequentially arranged from front to back, the brake structure is located between the low-pressure turbine (97) and the turbine rear bearing frame (98), and the low-pressure turbine (97) is a two-stage low-pressure turbine.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310560800.8A CN116357404A (en) | 2023-05-17 | 2023-05-17 | Braking structure for preventing turbine from flying after broken shaft of aero-engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310560800.8A CN116357404A (en) | 2023-05-17 | 2023-05-17 | Braking structure for preventing turbine from flying after broken shaft of aero-engine |
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| CN116357404A true CN116357404A (en) | 2023-06-30 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202310560800.8A Pending CN116357404A (en) | 2023-05-17 | 2023-05-17 | Braking structure for preventing turbine from flying after broken shaft of aero-engine |
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| CN (1) | CN116357404A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN118346391A (en) * | 2024-06-05 | 2024-07-16 | 中国航发湖南动力机械研究所 | Aviation turboshaft engine easy to maintain |
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2023
- 2023-05-17 CN CN202310560800.8A patent/CN116357404A/en active Pending
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN118346391A (en) * | 2024-06-05 | 2024-07-16 | 中国航发湖南动力机械研究所 | Aviation turboshaft engine easy to maintain |
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